1. Field of the Invention
The present invention relates generally to a gas turbine engine, and more specifically to a turbine rotor blade with tip peripheral cooling.
2. Description of the Related Art Including Information Disclosed Under 37 CFR 1.97 and 1.98
In a gas turbine engine, such as a large frame heavy-duty industrial gas turbine (IGT) engine, a hot gas stream generated in a combustor is passed through a turbine to produce mechanical work. The turbine includes one or more rows or stages of stator vanes and rotor blades that react with the hot gas stream in a progressively decreasing temperature. The efficiency of the turbine—and therefore the engine—can be increased by passing a higher temperature gas stream into the turbine. However, the turbine inlet temperature is limited to the material properties of the turbine, especially the first stage vanes and blades, and an amount of cooling capability for these first stage airfoils.
The first stage rotor blade and stator vanes are exposed to the highest gas stream temperatures, with the temperature gradually decreasing as the gas stream passes through the turbine stages. The first and second stage airfoils (blades and vanes) must be cooled by passing cooling air through internal cooling passages and discharging the cooling air through film cooling holes to provide a blanket layer of cooling air to protect the hot metal surface from the hot gas stream.
A turbine rotor blade rotates within a stationary shroud surface (referred to as a blade outer air seal or BOAS) in which a gap is formed between the blade tip and the shroud surface. Hot gas will leak across the blade tip gap due to a positive gap. This hot gas leakage typically over-heats the blade tip and reduces the blade life. The blade tip gap does not remain constant during engine operation due to factors such as different metal properties from the rotor and the blade and casing. The blade tip erosion due to an over-temperature and lack of adequate cooling is more so in the trailing edge region because of the thin airfoil walls. First stage turbine blades are exposed to the highest hot gas stream temperatures and thus the over-temperature problem is more of an issue.
FIG. 1 shows a prior art turbine blade with a three-pass serpentine flow circuit used to provide cooling for the blade. A first leg 11 provides cooling for a leading edge region while a third leg 13 provides cooling for the trailing edge region. The cooling air for the third leg 13 flows first through the first and second legs 11 and 12 where the cooling air is heated. The cooling air in the third leg 13 is mostly discharged out from a row of trailing edge cooling slots 15 with remaining cooling air being discharged out from a tip cooling hole 16 located in the trailing edge region. A tip cooling air hole 14 can also be used in the tip turn channel between the first and second legs 11 and 12 for the cooling of the blade tip and for producing a seal for the tip gap. FIG. 2 shows a flow diagram for the FIG. 1 blade. FIG. 3 shows a cross section top view for the cooling circuit of the FIG. 1 blade.